A gas turbine engine such as that used for powering an aircraft in flight includes a turbine rotor having a plurality of circumferentially spaced apart blades having tips spaced radially inwardly from a surrounding stator shroud for defining a predetermined clearance therebetween. The turbine extracts energy from hot combustion gases channeled between the blades which gases are exhausted from the engine to generate propulsive thrust. Gases which leak through the blade tip clearance decrease the efficiency of the engine and resulting thrust. Accordingly, the blade tip clearance is made as small as possible to minimize leakage therethrough without effecting undesirable tip rubs of the blades with the stator shrouds during operation.
Thrust droop is a conventionally known phenomenon which occurs during engine accelerations from low power to high power in which the thrust droops below its steady-state maximum level for up to several minutes. It is especially significant in cold engines wherein the turbine rotor and stator shroud have not reached steady-state temperature, and are still differentially expanding.
More specifically, during engine acceleration, the temperature of combustion gases increases which heats both the turbine rotor and stator shroud. The stator shroud has a fast thermal response and therefore increases in radius relatively quickly compared to the increase in diameter of the turbine rotor which has a relatively slow thermal response. Accordingly, the blade tip clearance transiently increases due to the greater expansion of the stator shroud over the turbine rotor which increases the leakage of the combustion gases therethrough leading to thrust droop. As the temperature of the turbine rotor reaches its steady-state value after the engine acceleration, the blade tip clearance will be reduced to its steady-state value which will eliminate thrust droop.
Accordingly, thrust droop is a transient occurrence whose effect varies from insignificant levels to substantial levels depending primarily on whether the engine is initially cold or hot. In a cold engine, which occurs immediately after engine startup and during idle or low power cruise operation, the turbine rotor and stator shroud are operated at their lowest temperatures and are therefore relatively cold. Upon occurrence of engine acceleration, thrust droop will occur since both the turbine rotor and stator shroud are significantly heated by the exhaust gases for increasing output power, with the stator expanding substantially faster than the rotor. In a hot engine, in contrast, the turbine rotor and stator shroud are already at an elevated temperature and, therefore, engine acceleration does not lead to significant temperature increase thereof. Accordingly, significant thrust droop does not occur. And, in a warm engine having a temperature between cold and hot, varying amounts of thrust droop will occur. Of course, the terms cold, warm, and hot are relative, but they are used herein to indicate the degree of blade tip clearance increase upon engine acceleration from maximum to intermediate to substantially no increase, respectively.
One method and apparatus for compensating for thrust droop in an aircraft gas turbine engine is disclosed in U.S. Pat. No. 4,581,889--Carpenter et al, entitled "Gas Turbine Engine Control," assigned to the present assignee. Carpenter et al disclose an augmented two spool engine having a variable area exhaust nozzle (VEN) in which thrust droop compensation is effected by selectively varying the exhaust are of the VEN to vary the turbine exit exhaust gas temperature designated T.sub.5. Carpenter et al disclose the use of two thermal models represented by LaPlace Transforms for expansion of the turbine rotor and stator shroud to infer the increase in blade tip clearance, and, therefore, infer thrust droop. The VEN is modulated in response thereto for selectively increasing the turbine exit temperature T.sub.5 to transiently increase thrust to compensate for thrust droop based on the thermal models. However, this thrust droop compensation method includes no provision for determining whether the engine is either hot, warm, or cold and, therefore, can reduce life of the turbine due to increase in turbine exit temperature T.sub.5 when not needed for thrust droop compensation.